Reverse core gear turbofan

ABSTRACT

A gas turbine engine has a fan at an axially outer location. The fan rotates about an axis of rotation. The fan delivers air into an outer bypass duct, and across a booster fan positioned radially inwardly of the outer bypass duct. The booster fan delivers air into a radially middle duct, and across a cold turbine into a radially inner core duct being directed into a compressor. From the compressor, air flows axially in a direction back toward the fan through a combustor section, and across an exhaust of the turbine section as directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method is also disclosed.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under Contract No.FA8650-09-D-2923/D013 awarded by the United States Air Force. TheGovernment has certain rights in this invention.

BACKGROUND

This application relates to a reverse core geared turbofan engine havinga turbine driven by fan air.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor section. The fan may also deliver air into abypass duct to provide propulsion. The air delivered into the compressoris compressed and moved into a combustion section where it is mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving them to rotate. The rotation of the turbinerotors in turn drives the fan and compressor sections.

Recently a speed reduction has been incorporated between a fan driveturbine and the fan. This allows the fan to rotate at a slower speedthan other components that may be driven by the same turbine. As anexample, a low or intermediate compressor is often driven by the fandrive turbine.

Another feature that has been incorporated into gas turbines is a“reverse core” engine. In a reverse core engine, the compressor deliversair in an axial direction toward a front of the aircraft and into acombustion section. The products of combustion pass downstream overturbine rotors, however, those turbine rotors are located in an axialdirection toward the front of the engine, and typically the fan.

SUMMARY

In a featured embodiment, a gas turbine engine has a fan at an axiallyouter location. The fan rotates about an axis of rotation. The fandelivers air into an outer bypass duct, and across a booster fanpositioned radially inwardly of the outer bypass duct. The booster fandelivers air into a radially middle duct, and across a cold turbine intoa radially inner core duct. Air from the inner core duct is directedinto a compressor, and then flows axially in a direction back toward thefan through a combustor section, and across a core turbine section. Airis then directed into the middle duct. A gear reduction drives the fanfrom a fan drive turbine section.

In another embodiment according to the previous embodiment, a shaftdownstream of the gear reduction relative to the fan drive turbinesection also drives the booster fan.

In another embodiment according to the previous embodiment, a shaftdownstream of the gear reduction is also connected to rotate with thecold turbine.

In another embodiment according to the previous embodiment, the fanbooster and the cold turbine rotate with a clutched shaft separate froma fan shaft driving the fan. A clutch selectively connects the clutchedshaft to the fan shaft such that the fan shaft can selectively drive theclutched shaft.

In another embodiment according to the previous embodiment, there are atleast a plurality of core turbine sections, with one of the plurality ofcore turbine sections driving the fan through the gear reduction and asecond of the core turbine sections driving the fan booster.

In another embodiment according to the previous embodiment, thecompressor section includes at least a first compressor section and asecond compressor section downstream of the first compressor section.The core turbine section includes at least a first core turbine sectionand a second core turbine section. The first core turbine section drivesthe second compressor section and the second core turbine section drivesthe first compressor section. The second turbine section and firstcompressor section operate at a slower speed and at lower pressures thanthe first turbine section and the second compressor section.

In another embodiment according to the previous embodiment, a first coldturbine section is positioned adjacent the booster fan. A second coldturbine section is positioned downstream of the first cold turbinesection in the path of air flowing through the inner core duct, andupstream of the compressor section.

In another embodiment according to the previous embodiment, one of thefirst and second cold turbine sections is provided with a flow diverterthat allows bypass of air around a rotor associated with one of the coldturbine sections.

In another embodiment according to the previous embodiment, a radiallyouter extent of blades associated with one of the cold turbine sectionsis spaced inwardly of a radially outer position for the flow diverter toallow bypass of air radially outwardly of the radially outermost extentof the blades of one of the cold turbine sections.

In another embodiment according to the previous embodiment, there are apair of flow diverters, being movable between a position allowing thebulk of the air delivered to the compressor section to bypass theturbine rotors by passing radially outwardly of the radially outermostextent of the blades, and the flow diverters being movable to analternative position wherein the great bulk of the air delivered acrossone of the cold turbine sections passes radially inwardly of theradially outermost extent of the fan turbine blades.

In another embodiment according to the previous embodiment, the flowdiverter is associated with the first cold turbine section.

In another embodiment according to the previous embodiment, the flowdiverter is associated with the second cold turbine section.

In another embodiment according to the previous embodiment, the coldturbine section is provided with a flow diverter that allows bypass ofair around a rotor associated with the cold turbine section.

In another embodiment according to the previous embodiment, a radiallyouter extent of blades associated with the cold turbine section isspaced inwardly of a radially outer position for the flow diverter toallow bypass of air radially outwardly of the radially outermost extentof the blades in the cold turbine section.

In another embodiment according to the previous embodiment, there are apair of flow diverters, being movable to a position allowing the bulk ofthe air delivered to the compressor section to bypass the turbine rotorby passing radially outwardly of the radially outermost extent of theblades, and being movable to an alternative position with a great bulkof the air delivered across one of the cold turbine sections passesradially inwardly of the radially outermost extent of the fan turbineblades.

In another embodiment according to the previous embodiment, the coldturbine section associated with the flow diverter is positioned adjacentto the booster fan.

In another embodiment according to the previous embodiment, the coldturbine section associated with the flow diverter is positioned at alocation adjacent to the compressor on an axial side of the compressorspaced away from the fan.

In another embodiment according to the previous embodiment, the coldturbine section associated with the flow diverter is positioned adjacentto the booster fan.

In another embodiment according to the previous embodiment, the coldturbine section associated with the flow diverter is positioned at alocation adjacent to the compressor on an axial side of the compressorspaced away from the fan.

In another embodiment according to the previous embodiment, the coldturbine section associated with the flow diverter is positioned adjacentto the booster fan.

In another featured embodiment, a method of operating a gas turbineengine includes having blades with a radially outermost extent, and aflow diverter that is operable to divert air radially inwardly of theradially outermost extent of the turbine blades, or allow air to passradially outwardly of the radially outermost extent of the turbineblades. The flow diverter is positioned to increase or decrease theamount of gas passing across that turbine rotor to increase or decreasea power output by said turbine rotor.

In another embodiment according to the previous embodiment, there are apair of flow diverters moving between a position which passes all of thegas radially inwardly of the radially outermost extent of the turbinerotors, and to a position which diverts the gas radially outwardly ofthe radially outermost extent.

In another embodiment according to the previous embodiment, the flowdiverter is moved to the position to allow bypass of the gases in lowpower conditions, such as when a gas turbine associated with the turbinesection is in a aircraft at cruise conditions, and the flow diverter ismoved to direct the air radially inwardly of the radially outermostextent at high power conditions, such as take-off for that aircraft.

These and other features of the invention would be better understoodfrom the following specifications and drawings, the following of whichis a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a first embodiment.

FIG. 2 shows a second embodiment.

FIG. 3 shows a third embodiment.

FIG. 4A shows a feature which may be incorporated into any one of theFIG. 1-3 embodiments.

FIG. 4B shows the first embodiment feature in a second operativeposition.

FIG. 5A shows another embodiment feature which can be incorporated intoany of the FIGS. 1-3 embodiments.

FIG. 5B shows the second embodiment feature in a second operativeposition.

FIG. 6A shows another embodiment feature which can be incorporated intoanyone of the FIGS. 1-3 embodiments.

FIG. 6B shows the third embodiment feature in a second mode of operativeposition.

FIG. 7A shows a fourth embodiment feature which can be incorporated intoany one of the FIGS. 1-3 embodiments.

FIG. 7B shows the fourth embodiment feature and a second operativeposition.

DETAILED DESCRIPTION

FIG. 1 shows a gas turbine engine 20 having a fan 22 delivering air intothree flowpaths, as an outer bypass propulsion flowpath 24, a middleflowpath 26 wherein the air will mix with exhaust from an exhaust duct54, and an inner flow duct 28 which will deliver air into a core inletduct 32 for the reverse core engine 20.

A fan booster 50 is positioned downstream of the fan 22 and furtherdrives the air into the flowpaths 26 and 28. A turbine 52 (or coldturbine) receives the air from the inner flowpath 28 and extracts energyfrom the air as it is driven to rotate.

The air from the turbine 52 passes into the inner core flowpath 28, theduct 32, and into a low pressure compressor 30. The air is compressedand delivered into a high pressure compressor 34. The air is mixed withfuel in a combustion section 36 and ignited.

Products of the combustion pass downstream over a high pressure turbine38, a low pressure turbine 40 and another low pressure turbine 44.Downstream of the low pressure turbine 44, exhaust gases exhaust fromthe duct 54, and into the middle airflow duct 26.

The turbine 40 drives a spool 42 to drive the low pressure compressor30. The high pressure turbine 38 drives a spool 39 to in turn drive ahigh pressure compressor 34.

The turbine 44 is a fan drive turbine, and drives a gear reduction 46 toin turn drive a shaft 48. The shaft 48 is operatively connected to drivethe fan blade 22, the fan booster 50 and the turbine 52. Notably, theturbine 52 may also extract energy from the air delivered by the fanbooster 50 to rotate the shaft 48.

The provision of a turbine driven by the “cold” air downstream of thefan booster 50 provides greater efficiency to the overall arrangement.

FIG. 2 shows another embodiment 120. In embodiment 120, components whichare generally the same as the FIG. 1 embodiment bear like numbers,however, increased by 100. Embodiment 120 differs in that the gearreduction 146 drives a shaft 147. The shaft 147 is clutched by clutch160 to a shaft 148 which drives the fan booster 150, and the turbine152.

The clutch may be engaged to provide greater efficiency by eithercapturing the rotation of the turbine 152, or allowing it to free rotateand drive the fan booster 150 on its own.

FIG. 3 shows yet another embodiment 220. A turbine 256 drives the gearreduction 260 to drive the fan rotor 222. Again, components which aresimilar to those in FIGS. 1 and 2 are identified by the same referencenumeral, only increased by 200.

A separate turbine 258 is connected to the cold turbine 262, and thebooster fan blade 250 by a spool 260. In this regard, the powerdelivered to the fan booster 250, and how the power generated by theturbine 262 is utilized, has some additional freedoms.

FIG. 3 also shows a second fan air turbine section 304 which ispositioned downstream of the duct 232 and leading into the inlet for thecompressor section 302. The turbine section 304 is operable to rotatewith a shaft 307 that rotates with the low pressure turbine 256 and thelow pressure compressor 302.

FIG. 4A shows another feature which can be incorporated into any one ofthe three above-referenced embodiments. As shown, the turbine blades 262have a relatively short radially outer edge 270. A flow diverter, whichcould be a bypass door 272 of some sort is shown in an operativeposition to increase power flow. This position may be utilized suchduring takeoff of an associated aircraft on a hot day.

Notably, while the features of FIGS. 4-7 are shown associated with theFIG. 3 embodiment, they would have application into any one of theembodiments illustrated in this application.

FIG. 4B shows another operative position wherein the door 272 is pivotedoutwardly to create a bypass flowpath 274 which avoids the blades in theturbine section 262. The bypass door could be opening during cruiseconditions. In either case, the bypass flow 274 is still directed intothe inner flow path 228, and to the inlet duct 232 (for example, shownin FIG. 3).

FIG. 5A shows a second embodiment feature wherein a pair of flowdiverters or doors 272 and 276 are utilized. During a cruise conditionas shown in FIG. 5A the door of 272 is pivoted outwardly as is the door276. Now, all air is diverted away from the turbine section 262 andthrough the bypass path 274.

FIG. 5B shows an alternative operative position where the doors 272 and276 are pivoted inwardly such that the great bulk of the air would nowbe directed across the turbine section 262. This position would beutilized during high power conditions such as a takeoff on a hot day.

FIG. 6A shows yet another embodiment 300. In embodiment 300, the turbine302 is positioned downstream of the inlet duct 232, and upstream of thelow pressure compressor 302. Door 308 operates similarly to the FIG.4A/4B embodiment to direct all air across the turbine blades in turbine304, and within the radial extent of the blades (that is, radiallyinwardly of the radially outermost extent 306).

FIG. 6A would be utilized in high power conditions such as duringtakeoff of an associated aircraft on a hot day.

FIG. 6B shows the door 308 pivoted radially outwardly to provide abypass flowpath 310 which is outward of the radially outer end 306 ofthe blades in the turbine 304. This would be utilized at low powerconditions such as cruise.

FIG. 7A shows another embodiment 350. Again, the turbine 304 ispositioned downstream of the duct 232, but upstream of the low pressurecompressor 302. A second door 354 is provided in addition to the door352. In the FIG. 7A position, the bypass 356 is opened, such as may beutilized during cruise conditions. The lower door 354 ensures that thegreat bulk of air avoids the turbine blades in turbine section 304.

FIG. 7B shows the position of the doors 352 and 254 during high powerconditions such as takeoff on a hot day. In this position, the greatbulk of the air is directed radially inward of the radially outermostextent 306 of the blades in the turbine section 304.

A schematic control 800 is illustrated in the figures and would operateto control the various components disclosed across this application. Thecontrol can be incorporated into a FADEC for the entire engine. A workerof ordinary skill in the art would be able to design such a controlgiven the teachings of this disclosure.

For purposes of this application, the terms “low” or “high” relative topressure or speed, and in core turbine and compressor sections simplyare to be taken as relative terms. That is, the “high” would rotate athigher pressures and typically higher speeds than would the “low,”although both might be at objectively high speeds and pressures. Inaddition, the term “cold” for the turbine sections downstream of thebooster fan simply imply they are not part of the core engine. They maywell operate at very high temperatures, even though they are referred toas “cold” in this application. The turbine sections which are in thecore engine could be called “core turbine sections” for purposes of thisapplication. The core turbine sections would typically be seeing highertemperature and pressure gases than would the “cold” turbine sections.

As should be understood, all of the gas turbine engines illustrated inall of these figures rotate about a central axis of rotation. Thefigures are generally illustrating only the upper half of that engine,and there is an axis of rotation shown generally in dashed line in eachof the figures.

A worker of ordinary skill in this art would recognize that manymodifications would come within the scope of this disclosure. For thatreason, the following claims should be studied to determine the truescope and content of this application.

What is claimed is:
 1. A gas turbine engine comprising: a fan at anaxially outer location, said fan rotating about an axis of rotation;said fan delivering air into an outer bypass duct, and across a boosterfan positioned radially inwardly of said outer bypass duct, said boosterfan delivering air into a radially middle duct, and said booster fandelivering air across a cold turbine into a radially inner core duct;air from said inner core duct being directed into a compressor, and thenflowing axially in a direction back toward said fan through a combustorsection, and across a core turbine section, and then being directed intosaid middle duct; and a gear reduction for driving said fan from a fandrive turbine section.
 2. The gas turbine engine as set forth in claim1, wherein a shaft downstream of said gear reduction relative to saidfan drive turbine section also drives said booster fan.
 3. The gasturbine engine as set forth in claim 2, wherein a shaft downstream ofsaid gear reduction is also connected to rotate with said cold turbine.4. The gas turbine engine a set forth in claim 3, wherein said fanbooster, and said cold turbine rotate with a clutched shaft separatefrom a fan shaft driving said fan, and a clutch selectively connectingsaid clutched shaft to said fan shaft such that said fan shaft canselectively drive said clutched shaft.
 5. A gas turbine engine as setforth in claim 1, wherein there are at least a plurality of core turbinesections, with one of said plurality of said core turbine sectionsdriving said fan through said gear reduction and a second of said coreturbine sections driving said fan booster.
 6. The gas turbine engine asset forth in claim 1, wherein said compressor section includes at leasta first compressor section and a second compressor section downstream ofsaid first compressor section, and said core turbine section includes atleast a first core turbine section and a second core turbine section,with said first core turbine section driving said second compressorsection and said second core turbine section driving said firstcompressor section, with said second turbine section and said firstcompressor section operating at a slower speed and at lower pressuresthan said first turbine section and said second compressor section. 7.The gas turbine engine as set forth in claim 1, wherein a first coldturbine section is positioned adjacent said booster fan, and a secondcold turbine section is positioned downstream of the first cold turbinesection in the path of air flowing through said inner core duct, andupstream of said compressor section.
 8. The gas turbine engine as setforth in claim 7, wherein one of said first and second cold turbinesections is provided with a flow diverter that allows bypass of airaround a rotor associated with said one of said cold turbine sections.9. The gas turbine engine as set forth in claim 8, wherein a radiallyouter extent of blades associated with said one of said cold turbinesections is spaced inwardly of a radially outer position for said flowdiverter to allow bypass of air radially outwardly of said radiallyoutermost extent of the blades of said one of said cold turbinesections.
 10. The gas turbine engine as set forth in claim 9, whereinthere are a pair of flow diverters, with said pair of flow divertersbeing movable between a position allowing the bulk of the air deliveredto the compressor section to bypass the turbine rotors by passingradially outwardly of the radially outermost extent of the blades, andsaid flow diverters being movable to an alternative position wherein thegreat bulk of the air delivered across said one of said cold turbinesections passes radially inwardly of the radially outermost extent ofthe fan turbine blades.
 11. The gas turbine engine as set forth in claim8, wherein said flow diverter is associated with said first cold turbinesection.
 12. The gas turbine engine as set forth in claim 8, whereinsaid flow diverter is associated with said second cold turbine section.13. The gas turbine engine as set forth in claim 1, wherein said coldturbine section is provided with a flow diverter that allows bypass ofair around a rotor associated with said cold turbine section.
 14. Thegas turbine engine as set forth in claim 13, wherein a radially outerextent of blades associated with said cold turbine section is spacedinwardly of a radially outer position for said flow diverter to allowbypass of air radially outwardly of said radially outermost extent ofthe blades in said cold turbine section.
 15. The gas turbine engine asset forth in claim 14, wherein there are a pair of flow diverters, withsaid pair of flow diverters being movable to a position allowing thebulk of the air delivered to the compressor section to bypass theturbine rotor by passing radially outwardly of the radially outermostextent of the blades, and said flow diverters being movable to analternative position with a great bulk of the air delivered across saidone of said cold turbine sections passes radially inwardly of theradially outermost extent of the fan turbine blades.
 16. The gas turbineengine as set forth in claim 15, wherein said cold turbine sectionassociated with said flow diverter is positioned adjacent to saidbooster fan.
 17. The gas turbine engine as set forth in claim 15,wherein said cold turbine section associated with said flow diverter ispositioned at a location adjacent to said compressor on an axial side ofsaid compressor spaced away from said fan.
 18. The gas turbine engine asset forth in claim 4, wherein said cold turbine section associated withsaid flow diverter is positioned adjacent to said booster fan.
 19. Thegas turbine engine as set forth in claim 13, wherein said cold turbinesection associated with said flow diverter is positioned at a locationadjacent to said compressor on an axial side of said compressor spacedaway from said fan.
 20. The gas turbine engine as set forth in claim 13,wherein said cold turbine section associated with said flow diverter ispositioned adjacent to said booster fan.
 21. A method of operating a gasturbine engine comprising the steps of: (a) operating a turbine sectionhaving blades with a radially outermost extent, and a flow diverter thatis operable to divert air radially inwardly of the radially outermostextent of the turbine blades, or allow air to pass radially outwardly ofthe radially outermost extent of the turbine blades; and (b) positioningsaid flow diverter to increase or decrease the amount of gas passingacross that turbine rotor to increase or decrease a power output by saidturbine rotor.
 22. The method as set forth in claim 21, wherein thereare a pair of flow diverters, with said flow diverters moving between aposition which passes all of the gas radially inwardly of the radiallyoutermost extent of the turbine rotors, and to a position which divertsthe gas radially outwardly of the radially outermost extent.
 23. Themethod as set forth in claim 21, wherein said flow diverter is moved tothe position to allow bypass of the gases in low power conditions, suchas when a gas turbine associated with the turbine section is in aaircraft at cruise conditions, and the flow diverter being moved todirect the air radially inwardly of the radially outermost extent athigh power conditions, such as take-off for that aircraft.